This invention relates to attitude control of spacecraft, and more particularly to a method for transfer from a spin-stabilized to a three-axis stabilized operating mode.
Artificial satellites are in widespread use for various purposes. For some purposes, as for example for communication purposes, a satellite may be required to direct an instrument such as an antenna toward a heavenly body, hereinafter referred to as "Earth".
The launching of an artificial satellite may be accomplished by a vehicle, such as a space shuttle, which places the satellite in a low Earth orbit. A satellite on a shuttle takes a path illustrated as 31 in FIG. 2. Path 31 extends from Earth 30 to a low Earth orbit 32, where the satellite is released. The satellite is often spin-stabilized when released from the shuttle, with a typical rotational speed of 10 rpm.
A perigee thruster or engine is operated at a point 34 of FIG. 2 to boost the satellite from low Earth orbit 32 into an elliptical transfer orbit illustrated as 36 in FIG. 2. The perigee of transfer orbit 36 is at the altitude of the low Earth orbit, and the apogee is at the desired final orbital altitude. For a geosynchronous satellite, the final orbital altitude is approximately 22,400 miles. The apogee location in FIG. 2 is occupied by spacecraft 8.
When the satellite arrives at the apogee of the transfer orbit, it is still spinning. The spin axis is often designed to be coincident with the thrust or .DELTA.V axis of the apogee kick thruster or engine. FIG. 1a illustrates a typical satellite 8 configuration, including a satellite body 10 spinning about an X axis 12, upon which an apogee kick thruster assembly 14 is centered. Stowed or undeployed solar panels illustrated as 16 and 17 face outward from the body of spacecraft 10, in a direction orthogonal to spin axis 12.
If the satellite is launched by an unmanned rocket rather than by the space shuttle, it may enter the transfer orbit directly, without an intermediary lower orbit. In either case, the satellite arrives at the apogee spinning about its X axis 12. Ideally, the satellite spin axis should be oriented so that the apogee kick thruster produces velocity to place the spacecraft in the orbital plane (the plane of FIG. 2), and to adjust the orbit size. When launch is from a space shuttle, the spin axis will be in the equatorial plane, whereas the unmanned rocket may often leave the spin axis inclined relative to the equatorial plane. If the spin axis is not in the equatorial plane, it is rotated into the equatorial plane, and the apogee kick thrusters are then fired in order to achieve the final circular orbit, 37 of FIG. 2.
When the circular orbit is achieved, the satellite is still spinning about its axis 12. Since the satellite is spinning, the solar panels cannot be deployed, and satellite maneuvers including attitude stabilization and general maintenance (operation) must be accomplished using battery power if the solar panels do not receive sunlight.
So long as X axis 12, the satellite spin axis, remains in the equatorial plane, the spin of satellite 8 allows sunlight to fall on solar panels 16 and 17 during its rotation and operating power can be maintained. The sun lies in the plane of the ecliptic, which is inclined (as much as 23.degree.) relative to the equatorial plane. Sunlight is always available in the equatorial plane, as illustrated by arrows 38 of FIG. 2, except at locations which are eclipsed by the Earth. The direction from which sunlight arrives depends upon the time of year. The eclipse is very short, and therefore the spinning satellite can remain active in the final orbit indefinitely. However, in order to be useful, certain satellites must be de-spun and must assume particular attitudes relative to the Earth. Two despinning and reorienting modes are in general use.
A first despinning and attitude stabilization system utilizes thrusters to almost completely despin the satellite about the X axis 12. An active three-axis control system must be engaged essentially immediately in order to avoid destabilization. The spacecraft is immediately re-spun to a low rate, such as 1/10 rpm, about an axis which is perpendicular to the boresight of a two-axis sun sensor. The two-axis sun sensor then scans across the sky under the influence of the spin. The sun sensor produces two-axis information defining the angle of the sun relative to the boresight. Once the sun has been acquired by the two-axis sun sensor, thrusters are used in a feedback manner to despin relative to the sun, so that the sun is on the boresight of the sun sensor. Additional thrusters are then used to spin up about the sun line to 1/10 rpm. An Earth sensor, oriented with its boresight perpendicular to the boresight of the sun sensor, then scans across the sky, looking for the Earth. However, the Earth can be acquired by the Earth sensor only twice per orbit, at 6 AM and 6 PM local time, at which time the sun and the Earth are at right angles to each other relative to the spacecraft. Once Earth and sun have been acquired, the satellite is de-spun relative to the Earth. The satellite has essentially no spin stabilization at this time. The momentum wheels are energized or spun up in order to transition from zero momentum to bias momentum control, and the solar arrays are deployed. This method has the advantage that it can be used with any spacecraft and it may be autonomous, but is disadvantageous in that the pointing accuracy depends upon the accuracy of the sun sensor, so high pointing accuracy requires an expensive high-accuracy two-axis sun sensor. Also, the spin-up of the momentum wheels and the deploying of the solar arrays must be accomplished quickly, because after the final despinning step, the lock of the sun sensor on the sun begins to be lost.
A second method for despinning the satellite and entering three-axis attitude control is the dual-spin turn. In the dual-spin turn, the spacecraft is first rotated by applying thruster torque so that x spin axis 12 is aligned with orbit normal (out of the page in FIG. 2). The spacecraft is equipped with a reaction wheel or momentum wheel illustrated in FIG. 1 as 18, with its spin axis 20 orthogonal to spin axis 12.
Wheel 18 is initially not spinning when the satellite enters its final orbit, such as the geosynchronous orbit illustrated as 37 of FIG. 2. The final attitude desired for the spacecraft is with the X axis directed toward the Earth, and with Z axis 20 aligned with the orbit normal. The dual spin turn maneuver is continued by applying a torque to spacecraft body 10 by accelerating reaction wheel or momentum wheel 18 of FIG. 1a. Acceleration of the wheel produces a torque about axis 20 tending to rotate body 10 of spacecraft 8, which rotation is resisted by the gyroscopic stiffness provided by the spacecraft spin about axis 12. This combination of forces results in a motion which directs the Z axis 20 out of the orbital plane, and which also rotates the X axis as illustrated in FIG. 1b, positioning it in the orbit plane, and leaving it spinning at a low rate. Since the momentum remains fixed, the spin of the spacecraft body is reduced as wheel 18 is accelerated. In order to maintain stability of the maneuver, some residual body spin must remain, and therefore the Z axis does not quite reach coincidence with orbit normal. In addition, the residual spin results in nutation of the spacecraft body.
By reference to FIG. 1b, with the Z axis substantially aligned with orbit normal, it can be seen that solar panels 16 and 17 receive no sunlight in any orientation, and therefore the satellite must operate using battery power until the completion of the maneuver including the deploying of solar panels 16 and 17. The batteries are designed to operate the satellite under full load (with all transponders operating) during eclipse, which is a relatively short period (less than an hour) during each orbit. Batteries are therefore capable of operating the satellite under light-load conditions for periods which typically range from two to ten hours. Thus, a race against time begins in the dual-spin turn maneuver, following the rotation of the Z axis into coincidence with orbit normal. This race requires that the nutation due to residual spin be damped, and the spacecraft despun. All motion of the spacecraft must cease before deployment of the solar panels before the batteries are depleted.
Damping of the nutation and despinning results in the desired final attitude. The damping may be provided by either passive or active means. For reliability purposes, passive damping is preferred. The passive damping is often provided by the sloshing of fuel in the fuel tanks of the spacecraft. The nutation damping time constant depends on mass properties and spacecraft geometry. Too long a time constant compared with electrical storage capacity may render the maneuver impossible.
An improved transition maneuver is desired.